JPH0820269B2 - Astronomical search and acquisition method for 3-axis stable spacecraft - Google Patents

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method for searching and capturing an celestial body of a three-axis stable spacecraft, and more particularly to a method for capturing the sun of a spacecraft utilizing a narrow field of view sun sensor.

[0002]

In conventional sun acquisition methods, spacecraft control systems typically use three sensors each having a square field of view of ± 30 ° to ± 60 °, which stops the rotation of the spacecraft. Such a wide field of view is needed because when the sun actually rotates, it must be wide enough so that the sun does not deviate from the field of view of the sensor in order not to stop immediately.

[0003]

A major problem with sensors having such a wide field of view is that such a wide field of view is not reflected by sunlight reflected from spacecraft accessories such as antennas or solar array panels. That is, the incident light causes a malfunction. An object of the present invention is to develop a method of capturing an celestial body such as the sun that can reliably capture an celestial body such as the sun by using a sensor with a narrow field of view that is not affected by such reflected light. is there.

[0004]

The present invention provides a celestial object search and acquisition method in a three-axis stable spacecraft having a first sensor and a second sensor each having a different field of view. To detect a celestial object within a predetermined angular range of a plane defined by first and second axes perpendicular to each other and a plane defined by the first and third axes in the three axes, respectively. Has a narrow field of view with a limited pointing direction,
About a second axis perpendicular to the first and third axes until it is within the field of view of the first sensor within the predetermined angular range of the plane defined by the axis of The field of view of the first sensor is rotated, and the angular position of the second axis when the celestial body is located within the field of view of the first sensor is initially set as the angular position of the second axis of the spacecraft. And then the angular position of this initialized second axis is corrected over time by calculation from the rotational speed measurement of the second axis so that the celestial body is defined by the first axis and the second axis. The position of the first sensor with respect to the sensor position detector when in the field of view of the first sensor within the predetermined angle portion of the plane, and the celestial body is defined by the third axis and the first axis. Located within the field of view of the second sensor within the predetermined angular range of the defined plane. The field of view of the second sensor is rotated about the third axis perpendicular to the first axis until the celestial body is located within the field of view of the second sensor, and the angular position of the third axis is initially set. The angular position of the third axis of the spacecraft is initialized as the angular position, and then the initialized angular position of the third axis is corrected over time by calculation from the rotational speed measurement of the third axis, A second object in which the celestial body is within the predetermined angular portion of the plane defined by the first axis and the third axis;
Detecting the position of the second sensor with respect to the sensor position detector when the sensor is in the field of view of the sensor.

Further objects, features and advantages of the present invention will become apparent from the appended claims, in conjunction with the discussion of the following description and the accompanying drawings.

[0006]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS The acquisition method of the present invention has been shown and described for sun acquisition on an earth orbit satellite, which is provided by way of example, and the method of the present invention may be used in orbit or through interstellar space. It can also be applied to the case where an celestial body of arbitrary brightness or darkness is captured by a moving spacecraft.

In the embodiment, (a) and (b) of FIG.
And three sun sensors, such as the sun sensor 10 shown in (c). The sensor 10 includes two solar cells 11 that output a small current when the sunlight is incident on the solar cells. The housing 12 encloses the solar cell and includes a pair of narrow slits 14 that define the field of view of the sensor. The narrow slit defines the field of view of the sensor. The narrow slit produces a narrow field of view in the direction transverse to the slit. By forming a narrow field of view in one direction, the chances of reflected light from surrounding equipment or the like entering the sensor is dramatically reduced, giving the spacecraft designer wide latitude in the location of the spacecraft sensor.

The prior art uses a sensor with a sufficiently wide field of view so that the spacecraft thrust engine will not stop the sun from being in the sensor's field of view even if the spacecraft rotation is temporarily stopped. are doing. For example ±
A square sensor field of view of 60 ° × ± 30 ° is used. The method of the present invention utilizes a very narrow sensor field of view and yet reliably captures the sun even if the incident line of the sun overshoots the sensor field of view due to sensor rotation. You can The narrower the field of view of the sensor, the greater the effect of avoiding reflected light from the spacecraft accessories and the like. The minimum angle must be maintained to ensure proper alignment of the field of view on the spacecraft. In the preferred embodiment, the angular width of the field of view in the narrow direction, shown as angle A in FIG.
It is as small as 5 °. However, the method of the invention can also be used for wider fields of view.

In the narrow direction indicated by arrow 13 in FIG. 1 (a) as the direction of rotation of the spacecraft, the sensor is
As long as the sun is within the sensor's field of view, it provides an analog signal that is approximately proportional to the rotation angle of the spacecraft. For sensor 10,
This proportional signal is 2 divided by the sum of the two outputs.
It is obtained as the difference between the outputs of the two solar cells. The signal indicating the presence of the sun is obtained by the sum of the two outputs. The sensor has a flat zero position corresponding to the position where sunlight is received equally by both solar cells, where the proportional output signal is zero.

Referring to FIG. 2, spacecraft 18 having spacecraft body 19 has three sensors 20, 30 and 40 similar to sensor 10 shown in FIG. The zero position of the sensor forms the sensing plane, shown as fan beams 21, 31 and 41, respectively. The spacecraft is a thrust engine 60 positioned to produce opposite torque about the yaw axis.
A and 60B, and thrust engines 62A, 62B, 62C and 62D located in the corners of the spacecraft for producing cross-coupled torques around roll and pitch axes
Is provided. An additional thrust engine is provided to provide spacecraft control redundancy as is normally done. Spacecraft controller 64 receives inputs from sensors 20, 30 and 40 for use in commanding a thrust engine for maneuvers described below. The controller also measures changes in spacecraft attitude.

As shown, the capture method shown in this embodiment is intended to maximize capture of the sun's line of incidence on roll axis 16. The roll axes x, 16 were chosen merely as an example to illustrate the present invention and are arbitrary such that they define three orthogonal axes with respect to the position of the spacecraft that indicate the desired position of the spacecraft towards the sun. It should be understood that the axes of can be selected. The actual position of the spacecraft sensor is not important. Only the direction of the field of view is important.

The sensor 20 has a boresight and is located on the spacecraft 18. The center of the wider angular range of the sensor field of view 22 is located on the roll axis, and the plane defined by the roll axis and the yaw axis. It has a zero position detection surface 21 arranged at. The other two sensors 30 and 40 have zero position detecting surfaces 31 and 41 in a plane defined by the roll axis and the pitch axis, but the center line of the wider angular width of the field of view is above and below the roll axis. Symmetrically wide angle range "B"
It is oriented at an angle "C" which is slightly smaller than half (Fig. 1). Therefore, each field of view 32 of sensors 30 and 40
And 42 certainly include roll axis 16. Alignment configurations for other axes are also possible by varying the search algorithm for the axis being searched. All three sensor fields of view include the axis last pointed to the sun,
The wide angle directions of view of 30 and 40 are the narrow angle directions of view of sensor 20. The fields of view 32 and 42 preferably include a range of ± 45 ° to ± 80 ° of a plane defined by the axis and the pitch axis.

The capturing method of the present invention is a unit sphere 50 of a spacecraft.
3 in relation to FIG. The initial conditions that are adaptive as acquisition methods include arbitrary attitudes, and gyro range and spacecraft body rotation speeds of a size allowed by the spacecraft structure. The first step is to zero the rotation speed of the spacecraft body around all axes to zero, and the period is allowed to reach near zero rotation speed. The threshold test at rotation speed is performed as a condition for proceeding to the next step. Suppose that the sun's line of incidence 53, which is incident on the spacecraft at the first sun position indicated by the sun 52, is incident on position 54 after the spacecraft body rotational speed is zeroed.

The next step is pitch acquisition, in which the rotational speeds of both the roll and yaw gyro axes are controlled to zero,
A scan rate of typically 0.75 ° per second is commanded to the pitch axis in a predetermined direction. It does not matter which direction the rotation is selected. Sensors 30 and 40
Monitors to detect the presence of the sun. Sun sensor
The position of the spacecraft's pitch axis angle, when detected at either sensing surface 31,41 of either 30 or 40, is set as the initial position with respect to the sun's line of incidence. That is, the pitch axis angle is initialized at this angular position. The angular position of the spacecraft's pitch axis is then modified over time by integrating the pitch gyrorate measurements.
The position and rotation speed control algorithm in controller 64 is energized at a pitch that returns the spacecraft to the zero position of the sun sensor, the sun's line of incidence is in the center of the field of view of sensors 30 and 40, and the sensor's proportional output is zero. And the presence of the sun is indicated. The proportional output signal from the solar cell allows the spacecraft to be held in the center of the narrow field of view regardless of the angular width of the field of view. This completes the sun capture about the pitch axis and ensures that the sun's line of incidence is aligned with the plane defined by the roll and pitch axes.

Since the spacecraft is controlled to return to the sun sensor zero position using an integral gyro reference value, the pitch angle may overshoot the narrow sun sensor field of view 32 or 42. There is no problem with. After the speed of rotation of the spacecraft body's pitch axis is zeroed, the proportional sun sensor pitch measurement is used to continue updating the pitch position, thereby avoiding gyro drift. During the above steps, the incident direction of the sun moves from the starting position 54 shown in FIG. 3 to the position 56 along the path shown. Since the spacecraft is rotating perpendicular to the narrow field of view of the sun sensors 30,40, pitch rotation cannot be stopped while the sun is in the sensor field of view. Overshoot of the sensor beyond position 56 occurs as shown by the dashed line to position 57, after which the spacecraft controller is aligned with the center of the field of view of sensor 30 in the direction in which the sun's line of incidence is narrow. Return the spacecraft to the zero position of sensor 30.

If the sun is not detected in the pitch during the above steps after sufficient time has been allotted for one complete revolution, this is because the sun first approaches either the positive or negative pitch axis and the sensor 30 and 40
Is not detected because it does not cover these polar regions with extreme plus and minus pitch axes. These areas are called "keyhole" areas. To compensate for the sun's initial position near the plus or minus pitch axis, the yaw angle adjustment is commanded to have an amplitude equal to the angular width of the keyhole plus a few degrees of margin. To be done. Good design practice keeps the keyhole at about ± 20 ° to avoid excessive energy consumption,
Keyholes of about ± 45 ° are easily tolerated. The yaw adjustment is sufficient to move the keyhole away from the sun so that the second attempt at the processing step ensures success in pitch sun capture.

A keyhole in the polar region of the pitch axis is needed to simplify the acquisition method. If the field of view of the sensors 30 and 40 includes or is close to the pitch axis, and if the direction of the sun is initially parallel to the pitch axis,
The sun can be detected by the sensor regardless of the position of the spacecraft. Different and more complex search algorithms are needed to achieve sun acquisition.

The sensor 30 or 40 in which the sun includes the pitch axis
After being obtained in the field of view of the sensor 30, it is held in the field of view of the sensor 30 or 40 by the pitch control law using the sun sensor attitude data from the sensor 30 or 40. On the other hand, the yaw axis scanning rotation speed of 0.25 ° to 0.75 ° per second is
It is commanded to provide the zero position sensing surface 21 of the sensor 20 facing the sun. The incident line of the sun is the position in Figure 3.
Take the aisle from 56 to position 58. The proper direction for the yaw scan is determined by observing whether the sun is available at sensor 30 or sensor 40 during the preceding step.

Inaccurate selections can occur if the sun is first detected by both sensors. This occurs when the sun is near the roll axis and is detected by both sensors 30 and 40 due to the overlap of their fields of view. In this case, the acquisition method simply selects and initiates a yaw scan. If the choice is correct, or if it does not scan one end of the sensor 30 or 40 indicating that the choice is incorrect, the sun will be captured in the yaw axis. If an incorrect selection is indicated, the yaw scan direction is reversed, causing the sensor 20 to scan towards the sun. The final capture of the yaw sun capture has a temporary overshoot of the field of view 22 of the sensor 20 and is performed in the same way as described for the pitch axis.

The logic for executing the acquisition method is that a command is transmitted to the spacecraft in real time from the controller 64 mounted on the spacecraft or from the ground based on telemetry. In the preferred embodiment, the full capture method is implemented onboard a spacecraft such that sun capture completes the voluntary initiation.

In the preferred embodiment, pitch scanning is stopped and yaw keyhole adjustment is performed. However, if the yaw adjustment amplitude is small, it may be performed during the scan about the pitch axis. One option, which simplifies the whole logic circuit, is to be perfected in a small part, i.e. less than a quarter of a revolution of the pitch scan, to move the area of the keyhole from the sun's line of sight in the direction of sight. The logic is to perform a portion of the yaw adjustment performed at each rotation of approximately the same pitch angle to ensure.

In the illustrated embodiment, the sensor 20 is used for the transfer of orbital motion and the direction is controlled for such motion. The wide angle field of view of the sensor 20 is required for transfer of orbital motion. Sensor 20 is used exclusively for sun capture,
Alternatively, if other methods are developed for the transfer of orbital motion, it is not necessary for the sensor 20 to have a wide angle view of the roll / yaw surface as sun capture on the roll / pitch surface is first achieved. A small field of view in both directions is needed that is large enough only to ensure alignment in the roll axis.

An advantage of the method of the present invention is that sun capture by a 3-axis stable spacecraft is accomplished using three sun sensors that have a relatively narrow field of view in one direction. This reduces the amount of reflected light entering the sun sensor and gives the spacecraft designer wide latitude in deploying spacecraft sensors with large reflective appendages such as antennas or solar cell arrays. This method utilizes a 3-axis gyro velocity and an integral velocity sensing with a 3-slit sun sensor that achieves sun capture with a narrow field of view sensor.

The invention is not limited to the actual arrangements or methods described above, but various changes and modifications may be made without departing from the scope of the invention as defined in the claims. It is understood. One such modification is the use of a solar sensor that uses the measured output from a single photo-sensing element that defines the sensing surface instead of the zero position defined by the output of a pair of solar cells. Is. Another possible modification is the use of a single sun sensor in the plane defined by the roll and pitch axes with any direction of rotation about the yaw axis used to reach the highest point of capture. . If this is possible, additional fuel is used to correct the first rotation about the wrongly oriented pitch axis y.

[Brief description of drawings]

1 is a front view, a plan view and an end view of a sun sensor used in a spacecraft for carrying out the method of the present invention.

FIG. 2 is a perspective view showing the orientation of the detection surface of the sun sensor attached to the spacecraft.

FIG. 3 is a spacecraft monosphere containing the field of view of the sun sensor used to demonstrate the method of the present invention.

[Explanation of symbols] 10,20,30,40… Sensor, 18… Spacecraft.

Front Page Continuation (72) Inventor Mike W. Trumasov, California 90603, USA 90603, Hoitier, Mikinda Court 15988 (72) Inventor Thomas Day Faber, USA 91367, Woodland Hills, Calvert Street 23346 (56) Reference JP-A-1-182719 (JP, A) JP-A-3-82912 (JP, A) JP-B 1-333400 (JP, B2)

Claims (1)

[Claims] 1. A celestial body search and acquisition method for a three-axis stable spacecraft having a first sensor and a second sensor having different fields of view , wherein the first sensor and the second sensor detect the celestial body.
For each of the first and second perpendicular to each other in said three axes
The plane defined by the second axis and the first and third
Pointing within a certain angular range of the plane defined by the axis of
Has a narrow field of view with limited direction, and the celestial body is in front of a plane defined by a first axis and a second axis.
Rotating the field of view of the first sensor about a second axis perpendicular to the first and third axes until it is within the field of view of the first sensor within a predetermined angular range; Second axis when the celestial body is in the field of view of the first sensor
With the angular position of as the first set angular position of the second spacecraft
The angular position of the axis is initialized, and then the initialized angular position of the second axis is set to the second axis.
According to the calculation from the rotation speed measurement of
Corrected, and the celestial body is in front of the plane defined by the first axis and the second axis.
When within the field of view of the first sensor within the given angle
Of the position of the first sensor detects a sensor position detector, heavenly bodies of the third axis and the plane defined by the first axis
A position within the field of view of the second sensor within the predetermined angular range.
Second axis centered on a third axis perpendicular to the first axis until placed
The third axis when the celestial body is positioned within the field of view of the second sensor by rotating the field of view of the second sensor
With the angular position of as the first set angular position
The angular position of the axis is initialized, and then the initialized angular position of the third axis is set to the third axis.
According to the calculation from the rotation speed measurement of
Corrected in front of the plane defined by the first and third axes
Within the field of view of the second sensor , which is within the predetermined angle
A method for searching and capturing an celestial body of a three-axis stable spacecraft, characterized by detecting the position of a second sensor related to the sensor position detector at the time .